A method for coating a sealing fin (2) on a component of a turbomachine, in particular on a blade tip (6) of a blade (1) of a turbomachine, with armoring (3, 30, 300), and to a corresponding component, in which method a blade (1) having at least one sealing fin (2) and a slurry which comprises particles of MCrAlY or particles for forming an MCrAlY layer (31), where M is nickel and/or cobalt, are provided, the slurry is applied onto the sealing fin and dried, and the sealing fin with the applied slurry is subjected to an aluminizing process so that the MCrAlY layer comprises an Al-rich sublayer (32).

A segment of an abradable ring seal for a turbomachine, includes: a reinforcing support; a composite fabric including fibers impregnated with polymerized resin, and a track of abradable material. The composite fabric is interposed between the reinforcing support 71 and the track of abradable material and is bonded to the reinforcing support, on the one side, and to the track of abradable material, on the other side. The segment of an abradable ring seal has a resistance to tearing in a direction perpendicular to the interface between the track of abradable material and the composite fabric, on the one hand, and between the composite fabric and the reinforcing support, on the other hand, greater than 5.10.sup.6 newtons per square meter. A process for the manufacture of such a segment is also described.

A coating system on a CMC substrate is provided, along with methods of its tape deposition onto a substrate. The coating system can include a bond coat on a surface of the CMC substrate; a first rare earth silicate coating on the bond coat; a first sacrificial coating of a first reinforced rare earth silicate matrix on the at least one rare earth silicate layer; a second rare earth silicate coating on the sacrificial coating; a second sacrificial coating of a second reinforced rare earth silicate matrix on the second rare earth silicate coating; a third rare earth silicate coating on the second sacrificial coating; and an outer layer on the third rare earth silicate coating. The first sacrificial coating and the second sacrificial coating have, independently, a thickness of about 4 mils to about 40 mils.

The present invention relates to a labyrinth seal for a turbine engine, in particular of an aircraft, comprising a rotor element and a stator element extending around the rotor element, the rotor element being suitable for rotating relative to the stator element about an axis of rotation having an axial direction (DA), the rotor element comprising an annular lip having an outer radial end extending towards an abradable element (57) carried by the stator element, the outer radial end of the annular lip having a corrugation in the axial direction (DA) and a non-zero axial expanse (E.sub.5) associated with the corrugation, the abradable element (57) comprising a plurality of cells (50a, 50b) arranged adjacent to one another along the axial direction (DA) and an ortho-radial direction (O), the cells (50a, 50b) comprising walls which extend in an essentially radial direction, the cells being distributed with a first cell density in a first densified annular zone (Z.sub.51) of the abradable element, said densified annular zone (Z.sub.51) being located opposite the radial end of the lip, said densified annular zone having an axial expanse less than or equal to the axial expanse of the outer radial end of the lip, the cells being distributed according to a reference density of cells outside said first zone, the first density being greater than the reference density.

A process of preventing spallation for a geometrically segmented thermally insulating top coat on an article, the process including forming a surface feature with a surface feature tip on a surface of the article; disposing the thermally insulating topcoat over the surface feature; and forming segmented portions that are separated by faults extending through the thermally insulating topcoat from the surface feature tip.

A turbine article includes a substrate with a geometric surface feature having a plurality of recesses recessed into the substrate. A ceramic topcoat is disposed over the geometric surface feature. The topcoat includes portions that are separated by faults extending through the topcoat from the geometric surface feature so the topcoat is segmented. A ceramic interlayer is disposed between the topcoat and the geometric surface feature.

A method of fabricating a blade includes fabricating an abrasive tip and then attaching the abrasive tip to a free end of an airfoil section of a blade. The abrasive tip comprises particles disposed in a matrix material.

A component for a gas turbine engine includes an airfoil section including a free end and an abrasive coating sprayed onto the free end, the abrasive coating including a polymer matrix and an abrasive filler, the abrasive filler between about 50%-75% by volume of the abrasive coating.